Ceramic matrix composite blade outer air seal

ABSTRACT

A blade outer air seal (BOAS) includes a tube of a ceramic matrix composite (CMC) material. A preform within the shell defines a mount for the BOAS. The preform is of a CMC material. A gas turbine engine and a method of forming a blade outer air seal (BOAS) are also disclosed.

CROSS-REFERENCE TO RELATED APPLICATIONS

This disclosure is a continuation of U.S. patent application Ser. No.15/686,906, filed on Aug. 25, 2017.

BACKGROUND

A gas turbine engine typically includes a fan section, a compressorsection, a combustor section and a turbine section. Air entering thecompressor section is compressed and delivered into the combustionsection where it is mixed with fuel and ignited to generate ahigh-energy exhaust gas flow. The high-energy exhaust gas flow expandsthrough the turbine section to drive the compressor and the fan section.The compressor section typically includes low and high pressurecompressors, and the turbine section includes low and high pressureturbines.

The compressor and turbine sections include alternating stages ofrotating blades and fixed vanes. The vanes direct flow at a desiredangle into the rotating blade stage. The rotating blade rows rotatewithin an engine case. A blade outer air seal is provided at eachrotating blade stage to establish an outer radial flow path boundary.Moreover, the blade outer air seal provides a clearance between a tip ofthe rotating blade stages and the outer radial flow path boundary.

Turbine engine manufacturers continue to seek improvements to engineperformance including improvements in engine assembly, materialcapabilities, and thermal, transfer and propulsive efficiencies.

SUMMARY

In a featured embodiment, a blade outer air seal (BOAS) includes a tubeof a ceramic matrix composite (CMC) material. A preform within the shelldefines a mount for the BOAS. The preform is of a CMC material.

In another embodiment according to the previous embodiment, the tube hasa first open end and a second open end at opposite sides of the BOAS,and the preform includes a first preform in the first open end facingcircumferentially outward and a second preform in the second open endfacing circumferentially outward opposite the first preform.

In another embodiment according to any of the previous embodiments, thefirst preform and the second preform define a curved surface defining afirst slot on the first end and a second slot on the second end.

In another embodiment according to any of the previous embodiments, thefirst preform and the second preform have primary fibers substantiallyfollowing a contour of a corresponding one of the first slot and thesecond slot.

In another embodiment according to any of the previous embodiments,includes at least one insert for each of the first preform and thesecond preform supporting a portion of the corresponding one of thefirst preform and the second preform.

In another embodiment according to any of the previous embodiments, thetube includes a substantially rectangular shape with a radially innersurface and a radially outer surface. The radially outer surfaceincludes a first cutout and a second cutout at respective first andsecond ends.

In another embodiment according to any of the previous embodiments, eachof the first end and the second end includes an end groove for a seal.

In another embodiment according to any of the previous embodiments, thetube has primary CMC fibers form one of a three-dimensional braid, aplurality of two-dimensional layers and a three-dimensional weave.

In another embodiment according to any of the previous embodiments, thetube has primary CMC fibers substantially following a longitudinallength of the BOAS.

In another featured embodiment, a gas turbine engine includes a case. Amount is attached to the case. A blade outer air seal (BOAS) has apreform disposed within a tube. The preform defines a slot for receivingthe mount. The tube and the preform are provided by a ceramic matrixcomposite (CMC) material.

In another embodiment according to any of the previous embodiments, thetube has a first open end and a second open end at opposite sides of theBOAS, and the preform includes a first preform defining a first slot inthe first open end facing circumferentially outward and a second preformdefines a second slot in the second open end facing circumferentiallyoutward opposite the first preform.

In another embodiment according to any of the previous embodiments, thefirst preform and the second preform have primary fibers substantiallyfollowing a contour of a corresponding one of the first slot and thesecond slot.

In another embodiment according to any of the previous embodiments,includes at least one insert for each of the first preform and thesecond preform supporting a portion of the corresponding one of thefirst preform and the second preform.

In another embodiment according to any of the previous embodiments, thetube includes a substantially rectangular shape with a radially innersurface and a radially outer surface. The radially outer surfaceincludes a first cutout and a second cutout at respective first andsecond ends.

In another embodiment according to any of the previous embodiments, eachof the first end and the second end includes an end groove for a seal.

In another featured embodiment, a method of forming a blade outer airseal (BOAS) includes forming a substantially rectangular tube of ceramicmatrix composite (CMC) material. A first preform and a second preform isformed from a CMC material. The first preform is assembled into a firstend of the tube and the second preform is assembled into a second end ofthe tube.

In another embodiment according to any of the previous embodiments, thefirst preform and the second preform are formed separate from the tubeto define a respective first slot and second slot and forming of thefirst preform and the second preform includes orientating primary fibersto substantially follow a contour of the respective first slot and thesecond slot.

In another embodiment according to any of the previous embodiments,forming the tube to include a radially outer surface and a radiallyinner surface and forming the radially outer surface to include a firstcutout at the first end and a second cutout at the second end.

In another embodiment according to any of the previous embodiments,assembling the first preform into the first end and the second preforminto the second end includes installing at least one insert forsupporting a portion of each of the first preform and the secondpreform.

Although the different examples have the specific components shown inthe illustrations, embodiments of this disclosure are not limited tothose particular combinations. It is possible to use some of thecomponents or features from one of the examples in combination withfeatures or components from another one of the examples.

These and other features disclosed herein can be best understood fromthe following specification and drawings, the following of which is abrief description.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 is a schematic view of an example gas turbine engine.

FIG. 2 is a schematic view of an example blade outer air seal.

FIG. 3 is a perspective view of an example blade outer air sealembodiment.

FIG. 4 is a cross-sectional view of an example blade outer air seal.

FIG. 5 is a schematic view of an example method of forming a blade BOAS.

FIG. 6A is a schematic view of an example primary fiber orientation.

FIG. 6B is a schematic view of another example primary fiber orientationfor the tube.

FIG. 6C is yet another example of the primary fiber orientation for thetube.

DETAILED DESCRIPTION

FIG. 1 schematically illustrates a gas turbine engine 20. The gasturbine engine 20 is disclosed herein as a two-spool turbofan thatgenerally incorporates a fan section 22, a compressor section 24, acombustor section 26 and a turbine section 28. Alternative engines mightinclude an augmentor section (not shown) among other systems orfeatures. The fan section 22 drives air along a bypass flow path B in abypass duct defined within a nacelle 18, while the compressor section 24drives air along a core flow path C for compression and communicationinto the combustor section 26 then expansion through the turbine section28. Although depicted as a two-spool turbofan gas turbine engine in thedisclosed non-limiting embodiment, it should be understood that theconcepts described herein are not limited to use with two-spoolturbofans as the teachings may be applied to other types of turbineengines including three-spool architectures.

The exemplary engine 20 generally includes a low speed spool 30 and ahigh speed spool 32 mounted for rotation about an engine centrallongitudinal axis A relative to an engine static structure 36 viaseveral bearing systems 38. It should be understood that various bearingsystems 38 at various locations may alternatively or additionally beprovided, and the location of bearing systems 38 may be varied asappropriate to the application.

The low speed spool 30 generally includes an inner shaft 40 thatinterconnects a fan 42, a first (or low) pressure compressor 44 and afirst (or low) pressure turbine 46. The inner shaft 40 is connected tothe fan 42 through a speed change mechanism, which in exemplary gasturbine engine 20 is illustrated as a geared architecture 48 to drivethe fan 42 at a lower speed than the low speed spool 30. The high speedspool 32 includes an outer shaft 50 that interconnects a second (orhigh) pressure compressor 52 and a second (or high) pressure turbine 54.A combustor 56 is arranged in exemplary gas turbine 20 between the highpressure compressor 52 and the high pressure turbine 54. A mid-turbineframe 58 of the engine static structure 36 is arranged generally betweenthe high pressure turbine 54 and the low pressure turbine 46. Themid-turbine frame 58 further supports bearing systems 38 in the turbinesection 28. The inner shaft 40 and the outer shaft 50 are concentric androtate via bearing systems 38 about the engine central longitudinal axisA which is collinear with their longitudinal axes.

The core airflow is compressed by the low pressure compressor 44 thenthe high pressure compressor 52, mixed and burned with fuel in thecombustor 56, then expanded over the high pressure turbine 54 and lowpressure turbine 46. The mid-turbine frame 58 includes airfoils 60 whichare in the core airflow path C. The turbines 46, 54 rotationally drivethe respective low speed spool 30 and high speed spool 32 in response tothe expansion. It will be appreciated that each of the positions of thefan section 22, compressor section 24, combustor section 26, turbinesection 28, and fan drive gear system 48 may be varied. For example,gear system 48 may be located aft of combustor section 26 or even aft ofturbine section 28, and fan section 22 may be positioned forward or aftof the location of gear system 48.

The engine 20 in one example is a high-bypass geared aircraft engine. Ina further example, the engine 20 bypass ratio is greater than about six(6), with an example embodiment being greater than about ten (10), thegeared architecture 48 is an epicyclic gear train, such as a planetarygear system or other gear system, with a gear reduction ratio of greaterthan about 2.3 and the low pressure turbine 46 has a pressure ratio thatis greater than about five. In one disclosed embodiment, the engine 20bypass ratio is greater than about ten (10:1), the fan diameter issignificantly larger than that of the low pressure compressor 44, andthe low pressure turbine 46 has a pressure ratio that is greater thanabout five 5:1. Low pressure turbine 46 pressure ratio is pressuremeasured prior to inlet of low pressure turbine 46 as related to thepressure at the outlet of the low pressure turbine 46 prior to anexhaust nozzle.

The geared architecture 48 may be an epicycle gear train, such as aplanetary gear system or other gear system, with a gear reduction ratioof greater than about 2.3:1. It should be understood, however, that theabove parameters are only exemplary of one embodiment of a gearedarchitecture engine and that the present invention is applicable toother gas turbine engines including direct drive turbofans, land basedturbine engines utilized for power generation as well as turbine enginesfor use in land based vehicles and naval propulsion.

A significant amount of thrust is provided by the bypass flow B due tothe high bypass ratio. The fan section 22 of the engine 20 is designedfor a particular flight condition—typically cruise at about 0.8 Mach andabout 35,000 feet (10,668 meters). The flight condition of 0.8 Mach and35,000 ft (10,668 meters), with the engine at its best fuelconsumption—also known as “bucket cruise Thrust Specific FuelConsumption (‘TSFC’)”—is the industry standard parameter of lbm of fuelbeing burned divided by lbf of thrust the engine produces at thatminimum point. “Low fan pressure ratio” is the pressure ratio across thefan blade alone, without a Fan Exit Guide Vane (“FEGV”) system. The lowfan pressure ratio as disclosed herein according to one non-limitingembodiment is less than about 1.45. “Low corrected fan tip speed” is theactual fan tip speed in ft/sec divided by an industry standardtemperature correction of [(Tram ° R)/(518.7° R)]^(0.5). The “Lowcorrected fan tip speed” as disclosed herein according to onenon-limiting embodiment is less than about 1150 ft/second (350.5meters/second).

The example gas turbine engine includes the fan 42 that comprises in onenon-limiting embodiment less than about twenty-six (26) fan blades. Inanother non-limiting embodiment, the fan section 22 includes less thanabout twenty (20) fan blades. Moreover, in one disclosed embodiment thelow pressure turbine 46 includes no more than about six (6) turbinerotors schematically indicated at 34. In another non-limiting exampleembodiment the low pressure turbine 46 includes about three (3) turbinerotors. A ratio between the number of fan blades 42 and the number oflow pressure turbine rotors is between about 3.3 and about 8.6. Theexample low pressure turbine 46 provides the driving power to rotate thefan section 22 and therefore the relationship between the number ofturbine rotors 34 in the low pressure turbine 46 and the number ofblades in the fan section 22 disclose an example gas turbine engine 20with increased power transfer efficiency.

Referring to FIG. 2 with continued reference to FIG. 1, a portion of theturbine section 28 is schematically illustrated and includes a turbinerotor blade 64 rotating relative to a radial surface 74 defined by aplurality of blade outer air seals (BOAS) 72. The turbine blade 64includes a tip 66 that rotates proximate to the radial surface 74defined by the BOAS 72. The example shown in FIG. 2 is of a single,rotating turbine blade stage and may also be utilized within thecompressor section 24.

The BOAS 72 are supported within an engine case 62 with a mount 70. Themount 70 may be an integral part of the case 62 or may be a separatepart attached to the case 62. A plurality of BOAS 72 form a full hoopcircumferentially about the engine axis A to surround the blades 64. TheBOAS 72 control leakage of core flow C in the gap 68 between the tips 66and the inner surface 74. The illustrated mount 70 is disposed betweeneach BOAS 72 and is one of a plurality of such mounts 70 disposed withinthe engine case 62. The gap 80 between each of the BOASs 72 is bridgedby a feather seal 78 that is assembled between adjacent BOASs 72.

The BOASs 72 encounter extreme pressures and temperatures and thereforeit is desirable to utilize materials that are capable of operating inthe harsh environments encountered within a gas turbine engine 20. Inthis disclosed example, each of the BOASs 72 are formed from a ceramicmatrix composite material. CMC materials include a plurality of fiberssuspended within a ceramic matrix. The fibers can, for example, beceramic fibers, silicon fibers, carbon fibers, and or metallic fibers.The ceramic matrix material can be any known ceramic material such assilicon carbide. The ceramic matrix composite material provides thedesired thermal capabilities to operate within the harsh environment ofthe turbine section 28.

Referring to FIGS. 3 and 4, the example BOAS 72 comprises a tube 82 withopen ends 94 into which are assembled and mounted preforms 84. In thisexample, the tube 82 comprises an open structure having a substantiallyrectangular shape in cross-section. However, the tube 82 may be shapeddifferently and remain within the contemplation of this disclosure. Thepreforms 84 define first and second slots 96 that correspond with ashape and contour of the mount 70. The example disclosed mount 70includes outward extending arms 75 that fit within the slots 96 definedby the preforms 84. Each of the open ends 94 also includes a groove 90for the feather seal 78.

The open ends 94 each have a corresponding cutout 88 within the topsurface 76 that is open the corresponding end 94. The cutouts 88corresponds with a profile of the mount 70 such that the mount 70 mayextend into the open ends 94 of the tube 82. The tube 82 may alsoinclude an opening 86 along the top surface 76 utilized to providecooling air or to reduce the total weight of the BOAS 72. One or severalopenings 86 may be utilized and are within the contemplation of thisdisclosure. The preform 84 is formed separate from the tube 82 andassembled into each of the open ends 94.

Referring to FIG. 4 with continued reference to FIG. 3, an interfacebetween adjacent BOASs 72 is shown with arms 75 of the mount 70extending into corresponding slots 96 defined by preforms 84 withinseparate BOASs 72. It should be appreciated that each BOAS 72 includesfirst and second open ends 94 that correspond with the mount 70 providedin the engine case 62. The preforms 84 are formed separately from thetube 82 and installed to define the slots 96 that correspond with themount 70.

Inserts 92 are provided along with each of the preforms 84 to supportcurved portions on a back side 85 of each preform 84. Each of thepreforms 84 is formed from a plurality of fibers that substantiallyfollow a contour of the desired slot 96. In this disclosed example, theslot 96 comprises a substantially c-shaped contour in cross-section thatcorresponds with arms 75 of the example mount 70. It should beappreciated that other shapes and contours could be utilized and arewithin the contemplation of this disclosure. Moreover, the specific fitbetween the preform 84 and the arms 75 of the mount 70 are such thatexcessive movement is prevented while accommodating relative thermalexpansion between the case 62, mount 70 and the BOASs 72. Additionally,each of the BOASs 72 is designed and dimensioned to accommodate thermalexpansion and movement relative to the rotating turbine blade 64.

Referring to FIG. 5 with continued reference to FIG. 4, a schematicillustration of a method of fabricating the example disclosed BOAS 72 isshown and generally indicated at 98. The method 98 includes an initialstep 100 of forming the tube 82. The tube 82 will include a width 122and a longitudinal length 124. A plurality of fibers will be includedalong the length 124 to provide the substantial structure of the tube82.

Referring to FIGS. 6A, 6B and 6C, the example tube 82 is formed from aplurality of fibers 120 disposed within a ceramic matrix that areprovided at a defined orientation. The desired orientation of the fibers120 can be one or a combination determined to provide the desiredmechanical properties of the tube 82.

In the example illustrated in FIG. 6A, the primary fibers 120 areorientated in a three dimensional braid.

Referring to FIG. 6B, the primary fibers 120 are layered in atwo-dimensional layers that extend substantially along the longitudinallength 124 of the tube 82.

Referring to FIG. 6C, a schematic view of another fiber orientation isillustrated and shows the primary fibers 120 orientated in asubstantially three-dimensional woven mat that extends in the directionof the longitudinal length of the tube 82.

Referring back to FIG. 5, formation of the preforms 84 is schematicallyillustrated at 102 and also includes formation of the inserts 94. Eachof the preforms are formed with primary fibers 116 suspended in aceramic matrix. The primary fibers 116 are orientated to follow acontour 118 that is utilized to define the slot 96. In this example, thefibers 116 substantially follow the contour 118 of the slot 96. Theinserts 94 are formed from randomly orientated fibers or othercompatible CMC material and fibers. The inserts 94 in this disclosedexample are formed separate from the preforms 84 and are shaped tocorrespond with the back side 85 contour of the preform 84. The inserts94 engage the back side 85 of the preform 84 to reduce and eliminate anyunsupported region or area once installed into the tube 82. Although,the example inserts 94 are disclosed as separate parts, formation of thepreform 84 to include integral structures on the back side 85 forsupport could be utilized and are within the contemplation of thisdisclosure.

The preforms 84 and inserts 92 can be formed using known CMC techniquesincluding layering of a number of CMC sheets, polymer infiltration(PIP), chemical vapor infiltration (CVI) and chemical vapor deposition(CVD). In these processes the primary fibers are provided as a preformthat is subsequently infiltrated with a ceramic matrix material. Byforming the tube 82, preform 84 and inserts 92 separately, theindividual structures have increased quality and can be formed withdensities and material properties that would be difficult to attain whenforming the entire BOAS as a single structure.

Each of the preforms 84 are initially formed in a larger size than isrequired to fit within the tube 82. As is shown at 104, each of thepreforms 84 is initially machined to provide a desired width 110. Theprimary fibers 116 are orientated to define the slot 96 and once thepreform 84 is initially machined to provide a desired width as indicatedat 110. The machining operation can include grinding, cutting or anyother machining operations understood to be compatible with CMCmaterials.

As indicated at 106, the height of the preform 84 is then adjusted tofit within the tube 82. All machining operations on the preform 84 aremade with respect to a datum schematically indicated at 114 thatcorresponds with the mount structure 70. It is the slot 96 of thepreform 84 that provides the origin to which all dimensions include thewidth 110, height 112 along with the shape of the slot 96 are orientatedsuch that the installed preform 84 corresponds with the features of themount 70.

Once the preform 84 is machined to the proper, desired size, it isinstalled within the tube 82 as schematically indicated at 108.Installation of the preform 84 into the end of the tube 82 defines theslots 96 within the completed BOAS 72. The preforms 84 are installedsuch that the slot 96 defined by the preform 84 corresponds with theopen ends 94 and cut out 88 of the tube 82.

Assembly of the preforms 84 and inserts 94 to the tube 82, 84 can beaccomplished by any means understood known by those skilled in the artfor adhesion of CMC materials to one another. In one example embodiment,the tube 82 and preform 84 are assembled in a partially cured conditionand then fully cured together to provide a desired adhesion andstructure. In another example embodiment a ceramic matrix material isfurther infused into a partially cured tube 82, preform 84 and inserts92 once assembled and finally cured to form one continuous structure.Moreover, other known processes and methods of joining CMC parts couldbe utilized within the contemplation of this disclosure.

Accordingly, the example BOAS 72 includes separately formed CMCcomponents to form different structures for mounting and definition ofthe boundary surface to increase build quality, strength and durability.

Although an example embodiment has been disclosed, a worker of ordinaryskill in this art would recognize that certain modifications would comewithin the scope of this disclosure. For that reason, the followingclaims should be studied to determine the scope and content of thisdisclosure.

What is claimed is:
 1. A blade outer air seal (BOAS) comprising: a tubeof a ceramic matrix composite (CMC) material; a preform within the shelldefining a mount for the BOAS, wherein the preform is of a CMC material.2. The BOAS as recited in claim 1, wherein the tube has a first open endand a second open end at opposite sides of the BOAS, and the preformcomprises a first preform in the first open end facing circumferentiallyoutward and a second preform in the second open end facingcircumferentially outward opposite the first preform.
 3. The BOAS asrecited in claim 2, wherein the first preform and the second preformdefine a curved surface defining a first slot on the first end and asecond slot on the second end.
 4. The BOAS as recited in claim 3,wherein the first preform and the second preform have primary fiberssubstantially following a contour of a corresponding one of the firstslot and the second slot.
 5. The BOAS as recited in claim 4, includingat least one insert for each of the first preform and the second preformsupporting a portion of the corresponding one of the first preform andthe second preform.
 6. The BOAS as recited in claim 2, wherein the tubecomprises a substantially rectangular shape with a radially innersurface and a radially outer surface, the radially outer surfaceincluding a first cutout and a second cutout at respective first andsecond ends.
 7. The BOAS as recited in claim 6, wherein each of thefirst end and the second end includes an end groove for a seal.
 8. TheBOAS as recited in claim 1, wherein the tube has primary CMC fibers formone of a three-dimensional braid, a plurality of two-dimensional layersand a three-dimensional weave.
 9. The BOAS as recited in claim 1,wherein the tube has primary CMC fibers substantially following alongitudinal length of the BOAS.
 10. A gas turbine engine comprising: acase; a mount attached to the case; a blade outer air seal (BOAS) havinga preform disposed within a tube, the preform defining a slot forreceiving the mount, wherein the tube and the preform are provided by aceramic matrix composite (CMC) material.
 11. The gas turbine engine asrecited in claim 10, wherein the tube has a first open end and a secondopen end at opposite sides of the BOAS, and the preform comprises afirst preform defining a first slot in the first open end facingcircumferentially outward and a second preform defines a second slot inthe second open end facing circumferentially outward opposite the firstpreform.
 12. The gas turbine engine as recited in claim 11, wherein thefirst preform and the second preform have primary fibers substantiallyfollowing a contour of a corresponding one of the first slot and thesecond slot.
 13. The gas turbine engine as recited in claim 12,including at least one insert for each of the first preform and thesecond preform supporting a portion of the corresponding one of thefirst preform and the second preform.
 14. The gas turbine engine asrecited in claim 12, wherein the tube comprises a substantiallyrectangular shape with a radially inner surface and a radially outersurface, the radially outer surface including a first cutout and asecond cutout at respective first and second ends.
 15. The BOAS asrecited in claim 12, wherein each of the first end and the second endincludes an end groove for a seal.
 16. A method of forming a blade outerair seal (BOAS) comprising: forming a substantially rectangular tube ofceramic matrix composite (CMC) material; forming a first preform and asecond preform from a CMC material; and assembling the first preforminto a first end of the tube and the second preform into a second end ofthe tube.
 17. The method as recited in claim 16, wherein the firstpreform and the second preform are formed separate from the tube todefine a respective first slot and second slot and forming of the firstpreform and the second preform includes orientating primary fibers tosubstantially follow a contour of the respective first slot and thesecond slot.
 18. The method as recited in claim 16, including formingthe tube to include a radially outer surface and a radially innersurface and forming the radially outer surface to include a first cutoutat the first end and a second cutout at the second end.
 19. The methodas recited in claim 16, wherein assembling the first preform into thefirst end and the second preform into the second end includes installingat least one insert for supporting a portion of each of the firstpreform and the second preform.